1. Field of the Invention
The present invention relates to a film cooling hole structure of a gas turbine moving blade in which arrangement of film cooling holes is optimized so as to enhance a cooling efficiency of the moving blade.
2. Description of the Prior Art
In a gas turbine moving blade known in the art, cooling air is flown in a serpentine cooling passage provided in the blade for effecting a convection cooling, and also cooling air is injected from film cooling holes onto a blade outer surface for effecting a film cooling.
FIGS. 6(a) and 6(b)are cross sectional views of one example of a gas turbine moving blade cooling structure in the prior art, wherein FIG. 6(a) shows an entire portion of the cooling structure and FIG. 6(b) shows a cross sectional view taken on line Bxe2x80x94B of FIG. 6(a). In FIG. 6(a), numeral 30 designates a moving blade, whose interior is sectioned by ribs 36, 37, 38, 39 to form a leading edge side cooling passage 31, a serpentine cooling passage comprising cooling passage portions 32, 33, 34 in a blade central portion, and a trailing edge side cooling passage 35, when the passage portions 32, 33, 34 communicate with each other in this order.
Cooling air represented by arrows 40 in a blade base portion enters the cooling passages, wherein the cooling air flowing in the leading edge side cooling passage 31 cools a blade leading edge portion and flows out of leading edge side holes as represented by arrows air 40a, the cooling air flowing in the cooling passage portions 32, 33, 34 cools the blade central portion and flows out of film cooling holes provided in a blade surface for effecting a film cooling of the blade surface as air represented by arrows 40b, and the cooling air flowing in the trailing edge side cooling passage 35 cools a blade trailing edge portion and flows out of a blade tip portion as represented by arrows air 40c and also flows out of a multiplicity of cooling holes provided in a blade trailing edge as air represented by arrows 40d. 
FIGS. 5(a) and 5(b) are cross sectional views of another example of a gas turbine moving blade cooling structure in the prior art, wherein FIG. 5(a) shows an entire portion of the cooling structure and FIG. 5(b) shows a cross sectional view taken on line Axe2x80x94A of FIG. 5(a). In FIG. 5(a), numeral 20 designates a moving blade, whose interior is sectioned to form a leading edge side cooling passage 21, a serpentine cooling passage comprising cooling passage portions 22, 23, 24, and a serpentine cooling passage comprising cooling passage portions 25, 26, 27 on a rear side-thereof, wherein the cooling passage portions 22, 23, 24 and 25, 26, 27 communicate with each other in this order, respectively.
Cooling air represented by arrows 41 in a blade base portion enters the cooling passages, wherein the cooling air entering passage (A) flows into the leading edge side cooling passage 21 and flows out of leading edge side holes as air represented by arrows 41a, the cooling air entering passage (B) flows into the cooling passage portion 22 to then flow through the cooling passage portions 23, 24 and flows out of film cooling holes provided in a blade tip portion as air represented by arrows 40b, and the cooling air entering passages (C), (D) flows into the cooling passage portion 25 to then flow through the cooling passage portions 26, 27 and flows out of a multiplicity of cooling holes of a blade trailing edge portion as air represented by arrows 41d. Thus, the blade is so constructed as to be cooled effectively in its entirety.
FIG. 4 is an enlarged explanatory view of portion x of FIG. 5(a) showing a film cooling hole structure in a cooling passage turning portion of the gas turbine moving blade in the prior art. The cooling passage portions 22, 23 are sectioned by a rib 51 and communicate with each other at a turning portion in the blade tip portion. In the blade tip portion, there are provided a multiplicity of film cooling holes 50. When the cooling air represented by arrow 41 flowing in the cooling passage portion 22 flows into the adjacent cooling passage portion 23 sectioned by the rib 51 as shown by arrow 41e, it does not flow along the rib 51 in the turning portion but separates therefrom as shown by arrow 41f, which results in causing a separation area 52 where a heat transfer rate is reduced. Further, as shown by arrow 41g, there arises a stagnation area 53 in a corner of the cooling passage portion 22, and the heat transfer rate is low in the stagnation area 53 also. Thus, there is caused a cooling non-uniformity in the cooling passage.
In the mentioned prior art gas turbine moving blades of FIGS. 5(a), 5(b), 6(a) and 6(b), there are provided the leading edge side cooling passage, the serpentine cooling passage of the blade central portion and the trailing edge side cooling passage and the cooling air is flown therethrough for blade cooling and the cooling air is also injected from the film cooling holes onto the blade outer surface for effecting a film cooling. However, the positions of the film cooling holes are not necessarily optimized, so that there arises the stagnation area of the cooling air in the cooling passage and also there is caused the separation phenomenon of the cooling air from the rib surface in the turning portion of the serpentine cooling passage. The stagnation area and separation area are areas where the heat transfer rate is reduced, thereby the cooling of the blade interior becomes non-uniform and this is one of the reasons for the cooling efficiency being reduced.
Thus, the present invention is made with a first object to provide a gas turbine moving blade cooling structure in which film cooling holes provided in a cooling passage are devised to be arranged so as to eliminate a stagnation area and a separation phenomenon of cooling air to thereby realize a uniform cooling in the cooling passage, and to enhance a cooling efficiency by eliminating an area where a heat transfer rate is low.
FIGS. 8(a) and 8(b) are cross sectional views of still another example of a gas turbine moving blade cooling structure in the prior art, wherein FIG. 8(a) shows an entire portion of the cooling structure and FIG. 8(b) shows a cross sectional view taken on line Bxe2x80x94B of FIG. 8(a). In FIG. 8(a), numeral 30 designates a moving blade, whose interior is sectioned by ribs 36, 37, 38, 39 to form a leading edge side cooling passage 31, a serpentine cooling passage comprising cooling passage portions 32, 33, 34 in a blade central portion, and a trailing edge side cooling passage 35, wherein the cooling passage portions 32, 33, 34 communicate with each other in this order. In each of these cooling passages, there are provided turbulators 48 for making a flow of cooling air therein turbulent to accelerate a convection to thereby enhance a heat transfer effect of the cooling air.
Cooling air represented by arrows 40 in a blade base portion enters the cooling passages, wherein the cooling air flowing in the leading edge side cooling passage 31 cools a blade leading edge portion and flows out of leading edge side holes as air represented by arrows 40a, the cooling air flowing in the cooling passage portions 32, 33, 34 cools the blade central portion and flows out of film cooling holes provided in a blade surface for effecting a film cooling of the blade surface as air represented by arrows 40b, and the cooling air flowing in the trailing edge side cooling passage 35 cools a blade trailing edge portion and flows out of a blade tip portion as air represented by arrows 40c and also flows out of a multiplicity of cooling holes provided in a blade trailing edge as air represented by arrows 40d. 
FIGS. 7(a) and 7(b) are cross sectional views of still another example of a gas turbine moving blade cooling structure in the prior art, wherein FIG. 7(a) shows an entire portion of the cooling structure and FIG. 7(b) shows a cross sectional view taken on line Axe2x80x94A of FIG. 7(a). In FIG. 7(a), numeral 20 designates a moving blade, whose interior is sectioned to form a leading edge side cooling passage 21, a serpentine cooling passage comprising cooling passage portions 22, 23, 24, and a serpentine cooling passage comprising cooling passage portions 25, 26, 27 on a rear side thereof, wherein the cooling passage portions 22, 23, 24 and 25, 26, 27 communicate with each other in this order, respectively. In this example also, like in the moving blade shown in FIGS. 8(a) and 8(b), there are provided turbulators 28 in each of the cooling passages so as to enhance a heat transfer effect of the cooling air.
cooling air represented by arrows 41 in a blade base portion enters the cooling passages, wherein the cooling air entering passage (A) flows into the leading edge side cooling passage 21 and flows out of leading edge side holes as air represented by arrows 41a, the cooling air entering passage (B) flows into the cooling passage portion 22 to then flow through the cooling passage portions 23, 24 and flows out of film cooling holes provided in a blade tip portion as air represented by arrows 41b, and the cooling air entering passages (C), (D) flows into the cooling passage portion 25 to then flow through the cooling passage portions 26, 27 and flows out of a multiplicity of cooling holes of a blade trailing edge as represented by arrows air 41d. Thus, the blade is so constructed as to be cooled effectively in its entirety.
In the mentioned prior art gas turbine moving blades of FIGS. 7(a), 7(b), 8(a) and 8(b) there are provided the leading edge side cooling passage, the serpentine cooling passage of the blade central portion and the trailing edge side cooling passage, wherein the turbulators are provided in each of the cooling passages, and the cooling air is flown therethrough for blade cooling and the cooling air is also injected from the film cooling holes onto the blade outer surface for effecting a film cooling. However, the positions of the film cooling holes are not necessarily optimized, so that there arises a separation area of the cooling air flow immediately after each of the turbulators in the cooling passage, and this separation area is an area where a heat transfer rate is reduced to thereby make the blade cooling non-uniform, which is one of the reasons for the cooling efficiency being reduced.
Thus, the present invention is made with a second object to provide a gas turbine moving blade cooling structure in which film cooling holes provided in cooling passages are devised to be arranged so as to eliminate a separation phenomenon of cooling air caused between each of turbulators, to thereby realize a uniform cooling in the cooling passages and to enhance a cooling efficiency by eliminating an area where a heat transfer rate is low.
In order to achieve the first object, the present invention provides the following.
A film cooling hole structure of a gas turbine moving blade is constructed such that an interior of the blade is sectioned by a rib into cooling passage portions communicating with each other so as to form a serpentine cooling passage. Cooling air for blade cooling is flown in the serpentine cooling passage to be flown out of the blade through film cooling holes. Two mutually adjacent cooling passage portions so sectioned by the rib are a cooling air flow upstream side passage and a cooling air flow downstream side passage. A portion of the film cooling holes is provided in an end corner portion of the cooling air flow upstream side, passage and a portion of said film cooling holes is provided at a position close to or in contact with a tip portion of the rib in the cooling air flow downstream side passage.
In the present invention, because a portion of the film cooling holes is provided in the end corner portion of the cooling passage portion on the cooling air flow upstream side of the two mutually adjacent cooling passage portions sectioned by the rib, the cooling air entering a stagnation area of the cooling air flow in this end corner portion flows outside of the blade through the film cooling holes provided thereat, so that cooling air flow occurs in the stagnation area and the heat transfer rate can be enhanced in the stagnation area in the end corner portion.
Further, there is formed a turning portion of the cooling air passage between the cooling air flow upstream side passage and the cooling air flow downstream side passage. In the cooling air flow downstream side passage, especially in the rib tip portion, the cooling air does not flow along the rib surface but separates therefrom, hence a separation area occurs in the rib tip portion and the cooling air flow therein becomes worse. Thus, in the present invention, in addition to the above-mentioned stagnation area, because a portion of the film cooling holes is provided in the separation are; that is, at the position close to or in contact with the rib tip portion, the cooling air flows outside of the blade through the film cooling holes provided thereat, so that cooling air flow occurs in the separation area and the heat transfer rate can be enhanced in the separation area.
The above-mentioned portion of the film cooling holes may be provided newly in the stagnation area and the separation area, or a portion of the film cooling holes provided conventionally may be moved to these areas, thereby such a low heat transfer area as the stagnation area or the separation area is eliminated and a uniform cooling of the moving blade and a longer life thereof can be attained.
Also, in order to achieve the second object, the present invention provides the following.
A film cooling hole structure of a gas turbine moving blade is constructed such that an interior of the blade is sectioned by a rib into cooling passage portions communicating with each other so as to form a serpentine cooling passage. Turbulators are provided on an inner wall of the serpentine cooling passage, and are arranged in multi-stages so as to cross a cooling air flow direction. cooling air for blade cooling is flown in the serpentine cooling passage to be flown out of the blade through a film cooling hole provided between each of the turbulators. The width of each of the turbulators is e and distance between a cooling air flow downstream side surface of each of the turbulators and a center of the film cooling hole downstream thereof is d. The film cooling hole positions between each of the turbulators are such that d/e is larger than 0 and smaller than 2 (0 less than d/e less than 2).
In the present invention, the film cooling hole is arranged so that d/e is larger than 0 and smaller than 2 (0 less than d/e less than 2); that is, the film cooling hole is provided close to or in contact with the rear side of the turbulator in the cooling air flow direction. Hence, a separation phenomenon of the cooling air flow wherein the cooling air is entrained reversely toward the rear side of the turbulator to separate from the wall surface can be eliminated. That is, because the film cooling hole is provided in a separation area, which is a low heat transfer area, caused by separation of the air flow in the vicinity of the rear side of the turbulator, the cooling air flows in the separation area to flow outside of the blade through the film cooling hole to accelerate a convection of the cooling air. Thus, the heat transfer rate is enhanced in the separation area and the cooling passage can be cooled uniformly.
Also, the present invention is made by combining the film cooling hole structure of the first object and the film cooling hole structure of the second object, to thereby provide a film cooling hole structure of a gas turbine moving blade which is able to achieve both of the first and second objects.